JP2012251542A - Variable cycle engine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a variable cycle engine including a low-bypass ratio turbofan to be used as a propeller of a small supersonic airplane, driven by a small engine, and comprising a simpler variable mechanism.SOLUTION: A tail cone installed in a core duct outlet immediately under a low-pressure turbine is allowed to move longitudinally. The low-pressure turbine expansion ratio is increased/decreased by changing the core duct outlet area when the movable tail cone moves longitudinally. The bypass ratio can be varied in response to a change in fan rotating speed. During takeoff, an exhaust speed is reduced and thrust is increased. During transonic speed increase and supersonic speed cruising, when the turbine inlet temperature and the whole pressure ratio are enhanced, an increase in the fan rotating speed is controlled to achieve high thrust.

Description

  The present invention relates to a small variable cycle engine (VCE) mounted on a small SST (Supersonic Transport) or SSBJ (Super Sonic Business Jet).

  In order to support expanding aviation demand, it is desired to develop a variety of air transportation systems by mutually complementing subsonic passenger aircraft for mass / low freight transportation and next-generation SST, which is high-speed / high-convenient transportation. In high-speed and highly convenient transportation, it is said that SSBJ or small SST, which has less environmental impact than large SST, is more likely to be realized prior to large SST.

  Since SSBJ goes straight between cities and arrives and departs from airfields near the city, airport noise reduction is the key to establishing SSBJ. On the other hand, in order to avoid the development risk, it is considered that the SSBJ engine is supported by an improved version of the current fighter engine. Therefore, it is necessary to make the bypass ratio (BPR) variable by simple modification, and the improvement must be based on a new concept different from the conventionally proposed method.

  Various types of variable cycle engines that make the BPR variable are considered, but the flow rate control for changing the BPR can be broadly divided into those that are performed on the compressor side and those that are performed on the turbine side. The BPR is changed by increasing or decreasing. When a turbine bypass engine is selected as a representative method for controlling the flow rate on the compressor side, one challenge for VCE becomes apparent. When air is extracted from the compressor outlet of the turbojet and reintroduced into the exhaust duct, it becomes a turbo fan. However, the total pressure difference between the extracted air and the core exhaust is too large, resulting in a large mixing loss. Therefore, it must be equipped with a variable area bypass injector or a diverter valve, which is too complicated for an improved small engine.

  Low-pressure turbine variable stator blades (LPT-VG: Low Pressure Turbine Variable Geometry) are used for the flow rate control on the turbine side, and a variable fan bypass injector (R--) is used to balance the pressure when the bypass flow and the core flow are mixed. There is VCE combined with VABI: Real-Variable Area Bypass Injector). This scheme reveals another challenge for VCE. The supersonic engine is hot at many locations, and the low pressure turbine (LPT) stationary blades must also be cooled, and lubricated at high temperatures to vary the angle of the stationary blades. This is an extremely difficult technology and unsuitable for improved small engines.

  Japanese Patent No. 3903270   Japanese Patent Application No. 2010-230042

  Journal of the Gas Turbine Society of Japan, Special Feature “Propulsion System for Supersonic Transporter (HYPR)”, Vol. 28, no. 1, January 2000   Journal of the Gas Turbine Society of Japan, Special Feature “Environmentally Compatible Next Generation Supersonic Propulsion System (ESPR)”, Vol. 32, no. 5, September 2004

  There are two problems to be solved. The first point is that LPT-VG is difficult to cool and lubricate, and the pressure adjustment mechanism is complicated in the mixing of the bypass flow and the core exhaust.

  Second, the conventional VCE changed the BPR by increasing or decreasing the core flow rate as described above. At take-off, the core flow rate was reduced to increase the BPR, and at supersonic speed, the core flow rate was returned to normal and the BPR was lowered to increase the specific thrust. However, today, when airport noise regulations are strict, a noise reduction device must be installed. For example, if a mixer ejector nozzle is used and external air is introduced and mixed with engine exhaust to reduce exhaust speed, a large thrust loss occurs. HYPR loses 7.5% thrust because it reduces the 15 dB noise at the mixer ejector nozzle. On the other hand, in the present invention, at the time of takeoff, by increasing the fan flow rate without reducing the core flow rate, the BPR is increased to lower the exhaust speed, and at the same time the thrust is increased. In addition, the supersonic speed is not based on the idea of lowering the BPR, but increases the total pressure ratio (OPR: Overall Pressure Ratio) to increase the thrust, and at the same time controls the LPT with simple means so that the fan does not overspeed. It is intended to establish a concept suitable for the VCE and to solve the problems of VCE.

  A conceptual diagram of the present invention is shown in FIG. In the figure, FAN is a fan, HPC is a compressor, CVSV is a compressor variable stator vane, CC is a combustor, HPT is a high pressure turbine, LPT is a low pressure turbine, TC may be a tail cone or a plug that can move back and forth, A7B is Bypass passage sectional area at engine position 7, A7C is a core duct outlet sectional area, RM is a lobe mixer, EN is an ejector nozzle. As shown in the figure, in order to make the present invention a simple variable mechanism that can be easily modified, the LPT-VG is eliminated and the tail cone TC disposed downstream at a lower temperature can be moved back and forth. When the core duct outlet area A7C is changed by the movement of TC, the LPT expansion ratio can be increased or decreased, and the fan speed can be controlled.

  The dimensionless air flow rate per unit time and unit area is defined by Equation 1 as a mass flux parameter (MFP).

Here, m is the mass flow rate, M is the Mach number, A is the cross-sectional area of the channel, Cp is the constant pressure specific heat, T is the total temperature, P is the total pressure, and κ is the specific heat ratio. If MFP (κ, M) is constant, the corrected flow rate increases from Equation 1 when the flow path cross-sectional area A is increased. When the state of the LPT inlet and the core duct outlet is expressed using MFP, m ₅ = m _7C from the law of conservation of mass, and thus Expression 2 is obtained.

When LPT is choked, A ₅ , MFP (M ₅ ) constant, left side P ₅ / P _7C and T ₅ / T _{7C in} Equation 2 are Mach number M _{7C at the} core duct outlet and area ratio A _7C / A ₅ It becomes a function. If the core duct outlet area A _7C is increased at a constant fuel flow rate, the core duct outlet Mach number M _7C decreases, and the MFP (M _7C ) decreases. However, the rate of increase in the area ratio A _7C / A ₅ is larger. Will increase. If the expansion ratio P ₅ / P _7C increases, the reciprocal T ₅ / T _{7C of the} temperature ratio also increases. However, since the denominator is the square root of the temperature, the rate of increase of the numerator increases.

In adiabatic flow, change in total enthalpy, it corresponds to the external work done during broaden the core duct exit area A _7C by the movement of the TC, Increasing the total enthalpy drop amount of LPT, donor from the working gas to the LPT Increased energy increases output. If the core duct outlet area A _7C just below the LPT is reduced, the decrease in the LPT back pressure can be suppressed, and the work of the LPT is reduced. Therefore, in this VCE, the number of fan rotations increases / decreases due to the TC moving back and forth, and the BPR changes.

When mixing the core duct outlet exhaust and the fan bypass flow at engine position 7, the core side static pressure p _7C and the bypass side static pressure p _7B must be equal as shown in Equation 3.

Accordingly, the Mach number M _7B of the bypass flow and the Mach number M _7C of the core flow at the position 7 are obtained by Expression 4.

Here, P is the total pressure, p is the static pressure, m = (κ−1) / κ, the suffix c is the compression side, and t is the turbine side. The ratio of the core duct outlet area A _7C and the bypass duct area A _7B is given by Equation 5.

Here, ρ is the density, V is the flow velocity, f is the fuel-air mixing ratio, and t is the static temperature. At the operating point satisfying both the expression 2 representing the relationship between the core duct outlet area A _7C and the LPT expansion ratio and the expression 5 immediately above, the core flow and the bypass flow can be mixed at the same static pressure, and the LPT expansion ratio can be changed at the same time. it can. That is, it can be said that the variable mechanism of the core duct outlet area A _7C by the movable TC has a function combining the LPT-VG and R-VABI of HYPR.

  The movable cone was already used for the Jumo004 in the early days of the jet engine. However, Jumo004 is not a turbofan but a turbojet, and its variable exhaust cone is not a mechanism that changes the speed ratio between the high pressure shaft and the low pressure shaft and changes the BPR. The variable exhaust cone of this engine is devised in order to keep the turbine inlet temperature (TIT) constant when controlling the thrust during flight. The movable cone of the present invention is installed not at the exhaust nozzle but at the core duct outlet, and by moving it back and forth, the speed ratio between the two axes is changed, and at the same time, the static pressure when mixing the bypass flow and the core flow is made equal. keep. Therefore, the technical idea of the movable cone of this engine and the variable exhaust cone of the turbojet are completely different.

  Table 1 shows the effect at takeoff and transonic speed obtained by increasing or decreasing the core duct exit area.

First, the effect at the time of takeoff is that when the core duct exit area A _7C is expanded to about 1.29 times the design point, the engine flow rate increases by 7.8%, the core flow rate increases by 2.4%, the BPR increases, and the exhaust jet speed increases. However, the thrust increases by 5.8% and the fuel consumption rate (sfc) decreases by 4.3%. Therefore, the thrust loss in the noise reduction device can be compensated. The problem in reducing jet noise is the ratio of noise reduction /% jet thrust loss. This target value for the ESPR mixer ejector nozzle is 4 dB. Therefore, if this target value can be achieved, this engine can reduce the product of 5.837% of the thrust increase and 4 dB, and the jet noise of 23.35 dB while maintaining the design point thrust.

Effect during transonic rise Next, when topped wall of the sound moves to supersonic flight, at elevated OPR enhance TIT, squeezed slightly to about 0.98 times the design point A _7C simultaneously fan When the increase in the relative correction speed is suppressed to about 110%, the increase in core flow rate (10.97%) is larger than the rate of increase in engine flow rate (6.0%), and at the same time the BPR decreases, thrust Will rise 18.8%. Therefore, high thrust exceeding the speed of sound can be ensured without an afterburner, so that fuel efficiency is greatly improved.

At the time of supersonic cruise, TIT becomes maximum temperature and squeeze A _7C in order to suppress the BPR which rises when a fixed cycle 0.91 times the design point, to keep the fan machine rotation speed to 100% Yes (see Table 2). This operation is possible because the STBJ having a flight Mach number of around 1.8 does not cause the CDT (compressor outlet temperature) to exceed the allowable material temperature due to the influence of the ram pressure.

  It is a conceptual diagram of this engine.   It is a fan operation map.   It is a compressor operation map.   It is a turbine characteristic figure at the time of takeoff.   It is a turbine characteristic figure at the time of a transonic rise.

  As a small engine for SSBJ, the LPT expansion ratio is changed by moving the tail cone back and forth with the aim of changing the BPR of the low bypass ratio turbofan with the simplest possible modification. This is achieved by controlling the fan speed (fan suction air flow rate) instead of increasing or decreasing the value.

  The fan characteristic map used for the cycle calculation performed as an example is shown in FIG. 2, and the compressor characteristic map is shown in FIG. Since maps and data of actual supersonic fighters are not available, both maps are virtual and not real maps.

In FIG, 1 is at takeoff (design point), 1-1 operating point that opened the core duct exit area A _7C during takeoff. 2 is the transonic speed of Mach 0.95, 2-1 is the operating point where the area A _7C is reduced by increasing the transonic speed (2 and 2-1 are indicated by white circles), 3 is the flight altitude 16 km, the flight Mach number 1.8 This is the operating point for the supersonic cruise. Table 2 shows the engine performance under the operating conditions shown in FIGS.

First, at the time of takeoff, the operating point is moved from the design point 1 to the operating point 1-1. The condition in this case is that the fuel flow rate is constant, and the high pressure turbine (HPT) and the LPT are in a choked state. Increasing the core duct exit area to open the A _7C total enthalpy drop amount increases expansion ratio of LPT, energy LPT fan speed and the engine flow rate increases increases because increasing.

As the LPT rotation speed increases, the core outlet condition of the fan changes, and at the compressor inlet, the core flow rate mc, the pressure P ₂ , and the temperature T ₂ also rise, and a higher flow rate of higher density flows into the compressor. . Therefore, the compressor correction rotational speed is slightly lowered. Fuel flow rate constant, HPT, chalk without increasing the work because conditions HPT both LPT, new thermodynamic equilibrium _{point, T 4} compared with immediately before widening the _{A 7C} is lowered but HPT inlet corrected flow does not change. Since the flow rate is pushed into the compressor on the core side of the fan, the pressure ratio increases, and a flow field in which BPR increases is formed. As a result, the thrust increases, sfc is reduced, and the jet speed Vj is also reduced (the exhaust speed in Table 2 is higher than the actual speed because cooling air or the like is not considered).

In transonic cruise 2, _A7C is slightly opened to maintain the LPT output in order to maintain the fan correction speed at 100%. Transonic increase 2-2 is a case where the engine flow rate is increased and the specific thrust is increased in order to break the sound wall and enter supersonic flight. Increase the fuel flow rate to increase the compressor pressure ratio and increase the thrust, but the fan speed will naturally increase if left as it is, so _A7C is throttled to prevent the LPT back pressure from decreasing and prevent the fan from over-rotating. . In supersonic cruise 3, TIT is the highest and becomes, kept for suppressing the BPR which rises when a fixed _cycle, the fan machine rpm squeeze _{A 7C} to 100%.

  Based on the cycle calculation, in order to analyze the LPT operability of this engine, the states of the LPT inlet and outlet (engine positions 5, 6, 7C) of 1 and 1-1 at takeoff were calculated. At takeoff 1 (design point), the LPT outlet 6 and the core duct outlet 7C have the same area. Table 3 shows the calculation results.

Interestingly, (1 1-1) both LPT outlet 6 axial Mach number _{M 6} of equal, there is little difference between the expansion ratio also. That is, the difference in expansion ratio between the two occurs in the core duct. A larger core duct exit area A _7C, the momentum flux is reduced, the working gas in the core duct is adiabatic expansion. Accordingly, the dynamic pressure of the working gas is reduced, and the kinetic energy is converted into LPT power through a change in momentum.

In general, the enlarged flow path (diffuser) has a reverse pressure gradient in a deceleration flow. Formula 2 shows that a flow field with a forward pressure gradient is formed in the decelerating flow when _A7C is expanded and expanded. This means that the core duct is not a stationary element. An equation expressing the total energy of the fluid is shown in Equation 6.

  From Equation 6, the total temperature of the fluid, and thus its total pressure, can only be changed by unsteady compression or expansion. Since steady flow is ∂ / ∂t = 0, it is not possible to add energy to the fluid or extract energy from the flow. Further, Equation 6 shows that the pressure (momentum flux) must be decreased to reduce the energy of the fluid as in the turbine, and is consistent with Table 3. The flow in the core duct becomes a steady flow and becomes a stationary element, that is, a diffuser, when the LPT is not choked.

The turbine characteristic diagram at take-off shown in FIG. 4 well represents the above physical phenomenon. P ₅ / P ₆ has no difference between 1 and 1-1 (both are choked), and only 1-1 has an increased expansion ratio at the core duct outlet 7C. FIG. 4 is consistent with the principle that the energy exchange between the rotating turbomachine and the fluid takes place only through an unsteady process.

  The characteristic curve at the turbine outlet is determined by the state quantity at the inlet and the power balance between the fan and the LPT. Therefore, when the volume flow rate is increased along this curve in the turbine choke state, the turbine back pressure is lowered from the constant inlet correction flow rate.

  FIG. 5 shows a turbine characteristic diagram when the transonic speed is increased. The increase in thrust at the transonic speed is only due to an increase in fuel. In this case, the role of the variable core duct exit area is to prevent an abnormal increase in the number of fan rotations as described above. The reaction degree R according to the temperature change amount is expressed by Equation 6.

  Table 4 shows the degree of reaction calculated from Table 3 using Equation 6.

When the area A _7C is expanded from Table 4, the reaction degree decreases (1-1), and when the A _7C is decreased, the reaction degree increases (2-1). From the definition of the reaction degree, it is understood that P ₆ = P _7C = 349.23 kPa after take-off 1-1. This is because there are no mechanical elements in the core duct flow path.

  In large SSTs, it is considered difficult at this stage to achieve the level of sonic boom strength at which land-based supersonic flight may be permitted. On the other hand, the small SST is premised on land supersonic flight. On the other hand, in recent years, regulations on airport noise by ICAO (International Civic Aviation Organization) have been increasingly strengthened, which is a big environmental issue for the next-generation supersonic aircraft, but the present invention makes it easy to create a low bypass engine capable of supersonic cruise. As a result of various modifications, the exhaust speed can be lowered at the time of takeoff and the thrust can be increased, so that the thrust loss in the noise reduction device can be compensated. The emergence of low noise at take-off and supersonic flying engines may create a new air transportation sector.

FAN: Fan f: Fuel / air mixing ratio HPC: Compressor CC: Combustor CVSV: Compressor variable stator blade HPT: High pressure turbine LPT: Low pressure turbine LPT-VG: Low pressure turbine variable stator blade TC: Tail cone that can be inserted and removed A: Channel cross-sectional area A7B: Bypass channel cross-sectional area at engine position 7 A7C: Core duct outlet cross-sectional area RM: Lobe mixer EN: Ejector nozzle 1: At takeoff (design point) 1-1: A7C open at takeoff 2: Transition Sonic cruise 2-2: Transonic increase A7C closed 3: Supersonic cruise A7C closed operating point subscript des: Design point, number: Represents the cross-sectional position of the engine

Claims (1)

  A movable tail cone that can be moved back and forth is provided at the core duct outlet directly under the low-pressure turbine of the low bypass ratio turbofan engine. At the same time as increasing / decreasing the turbine expansion ratio, it is possible to mix by equalizing the static pressure of the bypass flow and the core flow at the engine position 7, and at the time of takeoff, the core duct outlet area is widened to increase the fan rotation speed and the core flow rate. Increase the engine flow without reducing it, lower the exhaust speed and increase thrust, and at the time of transonic increase and supersonic cruise, increase the fuel flow and increase the thrust by increasing the total pressure ratio, the core duct exit area High thrust that enables supersonic flight by preventing the low-pressure turbine back pressure from falling and preventing the fan from over-rotating It can be obtained, a variable cycle engine comprising the variable mechanism of simple structure.

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